Gas turbine rotor cooling means



June 10, 1952 J. A. N. GALLIOT GAS TURBINE RoToR COOLING MEANS FiledFeb. 24, 1947 5 Sheets-Sheet June 10, 1952 J. A. N. GALLIOT 2,600,235

GAS TURBINE ROTOR COOLING MEANS Filed Feb. 24, 1947 3 Sheets-Sheet 2Inventor n /6 33 Ju LeslAnJrNorLert Gamm; BY @j m Attorneys June l0,1952 J. A. N. GALLloT 2,600,235

GAS TURBINE ROTOR COOLING MEANS Filed Feb. 24, 1947 3 Sheets-Sheet 5ATTORNEYS Patented June 10, 1952 UNI TED STAT E S PATE NT OFF ICE GASTRBNE ROTOR COOLING S JlilesAndr Norbert Galliot,'London,-England`Application February 24, 1947, Serial No. 730,446 In Great BritainFebruary 25, 1946 (Cl. (S-35.6)

13Claims. 1

This invention relates to lgas turbines such as are vemployed for*aircraft propulsion and like purposes, and comprising for exampleaperipherally bladed Vrotor `secured 'upon a shaft which drives acompressor feedingair to one or lmore combustion `chambers furnishing7hot gases to drive the peripheral blades Aof the rotor.

The main object of kthe 'present invention is to provide animproved gasturbine` so arranged and constructed that its rotor affords passage forcold air flowing tothe compressor intake whereby the rotor iseffectively cooled, ltogether with its blades, shaft and bearings.

A secondary vobject of the invention is to provide an vimproved gasturbine Yhaving a rotor fitted With vexternal blades driven by -hotgases and "internalvanes forcing cold air axially to a compressorfeeding a combustion chamber which furnishes the hot gases, said bladesand vanes being `:ln thermally conductive relation rto trans- `mit heatfrom said blades to said vanes vfor dissipation inthe `cold air passingybetween said vanes.

A further object of vthe invention is to -provide an improved gasturbine having anV apertured rotor which constitutes a primary elementof an auxiliary compressor forcing the air `admitted through 'the rotorapertures Yto the lintake of rthe main compressordriven by theturbineshaft.

Another object of the invention is to prov-ide an improved gas turbine'having' an apertured rotor 'and an air compressor mounted towards'opposite ends of a Adriving shaft, with a tubular casing around theshaft, the airadmitt'ed through the rotor apertures flowing `along theinterior of 'the casing to the intake of the-compressor.

A still further `object of the invention is to provide an improved gasturbine having an apertured and externally bladed rotor and a maincompressor mounted towards opposite ends of a driving shaft, with atubular casing around :the shaft-an axial-typecompressor inside thecasing, and an Aannular series of combustion chambers spaced around thecasing, the air' Vadmitted through the rotor apertures being forcedalong rthe interior-of the casing `.through the axial-type compressor tothe intake ofthe main compressor which delivers `compressed air v'to thecombustion chambers, and the latter furnishing hot gases to drive theexternal blades of the rotor.

A specific object ofthe invention is `to lprovide an improved gasturbine for jet-propelled aircraft, with or Vwithout an airscrew drivenby the turbine shaft, in which an apertured turbine rotor and its aircompressor 'are mounted towards oppositeends of the turbine shaft, withanannular seriesof combustion chambers-spaced around the shaft, the airadmitted through the rotor apertures flowing to the intake of thecompressor which 'delivers compressed air .to wthe combustion chambers,the latter furnishing hot gases to drive external blades upon ytheturbine rotor, and the gases `exhausted from the turbine `beingcollected to form a propulsive jet.

Other objects and advantages of the invention will hereinafter appearfrom the following description, given `with reference tothe accompanyingdrawings,in which:

Fig. 1 is a half sectional elevation of the turbine;

Fig. 2 isa partial development of the rotor on a larger scale, showingsix of the turbine blades;

Fig.3 is a partial end-view of the rotor, showing six of the externalturbine blades as well as the corresponding internal vanes forming theprimaryelement of an axial-type compressor;

Fig. 4 is a detail on a larger scale, showing a modified form ofplatform;

Fig. 5 is a partial end-view of the rotor hub mounted upon the turbineshaft, showing the method of ,assembling the roots ofthe compressorvanes;

Fig. 6 is a `half-section of the rotor hub, seen at right angles to Fig.5 takenat the center line;

Figs. 7 and y8 are part-sectional views, showing an exhaust manifold forfeeding a propulsiva-jet for aircraft;

Fig. 9 is a half sectional elevation of a modified turbine arranged todrive an airscrew for aircraft propulsion and also to deliver a jet forthe same purpose; and

Fig. 10 is a half sectional elevation of another modiilcation'of turbinearranged for pure jet propulsion of an aircraft.

In the construction illustrated in Fig. l, the turbine shaft I0, drivenby a rotor I I at the forward end, drives an air compressor I2 mountedtowards theother end ofthe shaft; this compressor feeds a number ofcombustion chambers I3 spaced around the shaft, each chamber beingprovided with fuel-injection means I3a and delivering the hot'gases ofcombustion through guide vanes I4 to the external blades I5 of theturbine rotor.

The central portion of `the rotor II is apertured by providing it withspokes or webs which are shaped as helicoidal vanes I6 (see Figs. 2, 3,5 and 6) to constitute a prima-ry element of an axial-type compressor;the cold air may bev admitted directly to the turbine rotor II by aflared inlet, opening in the direction of forward flight.

A shaft bearing, which is preferably of the selfaligningthrust-resisting type, is carried by radial arms or webs I8 inside theturbine stator, these arms being preferably formed as guide vanes toco-operate with the helicoidal vanes I6 of the rotor; in order toprotect the rotor against the entrance of flying bodies, a wire net I9amay be supported by brackets or arms I3 connected at their outer ends tothe exhaust manifold and at the inner end to a hub enclosing a bearingI1 for the turbine shaft.

The air compressor I2 may be of any suitable type, for examplecentrifugal; it is preceded by an axial-type compressor with multiplerotors 2G and stators 2I inside an induction tube or shaft casing 22, towhich the cold air is fed by the primary element consisting of theinternal rotor varies I6 already mentioned. The tube or casing 22, ofdiameter substantially equal at one end to that of the central portionof the rotor and at the other end to that of the compressor l2,surrounds the axial-type compressorl 20, 2I upon the shaft IIIconnecting the turbine rotor to the air compressor at the rear, thistube or casing being preferably made of or lined with thermallyinsulating material, as indicated at 24, to prevent undue heating of theincoming air by the combustion chambers I3 outside the tube or casing.

The centrifugal compressor I2 may be provided with two non-symmetricalworking faces, the front one operating as main compressor in tandem withthe axial compressor which delivers to it, and the rear face operatingin parallel as a separate source of compressed air, taking in airthrough suitable openings 213 at the rear end of the compressor casing;air may also be drawn through holes 21 in the shaft I0, adjacent to thisrear face 25, the shaft being made hollow to admit air at one or bothends. The separately acting rear compressor 25 will of course besuitably dimensioned to give the desired capacity; the two deliveries ofcompressed air from the front and rear faces may be merged together atthe periphery of the double-faced compressor I2, 25 before admission tothe combustion chambers I3. As illustrated in lFig. l, the compressorI2, 25 delivers into a manifold 28 having separate branches 28a leadingto the respective combustion chambers, suitable baffles or deflectors 23being provided inside the manifold to guide the compressed air into thebreech ends of tie chambers.

The one or more compressors may feed an nular combustion chamber,instead of the several separate combustion chambers I3 generallyutilized, the chamber or chambers in the respective cases furnishing hotgases to drive the turbine rotor II at the forward end 0f the shaft I0.

The rear bearing 30 of the turbine shaft may be mounted in a plate 3|supported by brackets 32 from the body or casing of the compressor I2,beyond which the shaft extends to operate auxiliary equipment (notshown) such as mechanism for feeding the fuel to the combustion chamberor chambers, or any other devices that may be desired for the operationof an aircraft or for other purposes.

Since the primary compressor vanes I6 are effectively cooled by theincoming air, the central turbine blades I5 to the cold vanes IB, theirjunctions are preferably in metallic continuity, for

example, they may be formed by segmental plates or platforms 33 integralwith the blades and vanes. These platforms are assembled to constitute acomplete annulus, of which a portion is shown in Figs. 2 and 3, theopposite edges of the platforms being engaged by clamping rings 34 ofchannel or V-section, held together by bolts 35 threaded into screwednipples 36 of similar design to those employed on cycle wheel spokes. Asseen in Fig. 1, the annulus formed by the assembled plates or platforms33, with the clamping rings 34, fits into grooves 41 on the adjacentedges of the stator walls; further, on the side directed towards theexhaust, a thin lip 48 projects from the roots of the turbine blades I5so as to overlap the grooved edge of the stator wall on this side. Dueto the high velocity of the gases acting upon the turbine blades,leakage of the gases into the interior of the rotor, which would polluteand heat up the incoming air, is prevented by the frictionless glandsprovided by these means.

Each of the segmental plates or platforms 33 may be made integral withone turbine blade I5 and the corresponding vane I6, in order to securethe greatest possible conduction of heat between those parts; due to thecontrary curvature of the blade and vane, as seen in Fig. 2, it may beneces sary to give the platforms a rhomboid or nonrectangular shape, asshown in that figure.

Fig. 4 illustrates a modified form of platform 33 made integral with therespective turbine blade I5 and rotor vane I6, in which the obliquesides of the rhomboidal platform are castellated or provided withalternate projections and notches to interengage with the sides of theadjacent platforms; the ends of the platform may be made with squarecorners, as shown, and notched for lightness.

The inner ends or roots of the rotor spokes or vanes are mounted uponthe shaft I0 by forming these roots 31 to a wedge-shape, as shown inFig. 5, the angle between the opposite faces of each wedge being equalto the angular spacing of the vanes; for example, with seventy-twoVanes, the included angle of each wedge will be five degrees. The wedgesare arranged to engage at their inner ends with splines 38 upon theexterior of the shaft; as shown in Fig. 5, the splines are each of anangular extent of five degrees, the inner ends of alternate wedgesresting upon the outer faces of the splines and those of theintermediate ones being made slightly longer in the radial direction soas to engage in the spaces between the splines, thus securing positivetransmission of the rotary motion. The wedge-shaped roots 31 are clampedtogether by means of hub-plates 39; each of these plates is providedwith circumferential grooves 40 on its inner face, the grooves fittingupon arcuate ribs 4I on the opposite faces of the wedges, and the platesbeing tightly held in engagement with the wedges, for example bypressure exerted along the shaft by a nut 42 securing the inner race ofthe bearing I1 (Fig. 1).

, Theinvention procures a considerable economy of space and dimensions;the tubular casing 22, into the open front end of which the air entersby ram effect, with initial compression by the primary elementconstituted by the vanes I3 of the apertured rotor and furthercompression by the axial-type compressor 2li, 2I inside the casing 22,is made convergent towards the main compressor I2 at the rear end, thismain compressor being of relatively small diameter and capacity becauseof the reduction of Volume of the air corresponding to its increasedpressure due to the initial and `secondary compressions. Theinventionalso .furnishes a solution of the problem of cooling the turbine rotor(including the blades I5 exposed to the combustion gases), the drivingshaft I0 and `its bearings; it follows that at least the central portionof the turbine rotor can be constructed of steelor other material chosensolely from the point of View of its mechanical qualities, withoutregard to the questions usually involved by high temperature-ofoperation. Where, however, some pre-heating of the air is desired, thethermal insulation of thecasing 22 (such as indicated at 24) may bedispensed with, wholly or in part, in order to facilitate the transferof heat from the combustion chamber or chambers I3.

The gases exhausted from the turbine may be collected in amanifold ofannular form having one or more lateral outlets for discharging thegases rearwardly; Figs. 7 and 8 illustrate a manifold 43 fitted to thestator of the turbine, having ltwo semi-annular branches of spiral shapeleading to two outlets 44 which are united by a Y- pipe fitting 45 todeliver a single jet for the propulsion of the aircraft or the like.Alternatively, the exhaust gases may be led to suitable reactiondevices, such devices being utilized for the propulsion of the aircraftor the like, as described in my prior U. S. Patent No. 2,224,260, datedDecember 10, 1940; an example of such reaction device, in which thegases are reversed through substantially 180 degrees by a deflectingsurface, is hereafter described with reference to Fig. 9.

The invention may be applied to gas turbines in general, apart fromtheir employment for jetpropelled aircraft.

Fig. 9 illustrates the arrangement of a turbine differing slightly fromthat shown in Fig. l, and having the forward end of the turbine shaft H0fitted with suitable speed-reducing gear IOI for the driving of anairscrew |02, the air flowing from the screw entering with ram effectinto the apertured central portion of the rotor III; the rotor isprovided with internal vanes I It of helicoidal form, cooperating withstationary guide vanes IIB and constituting the primary element of anaxial-type compressor with multiple rotors and stators I2I, insideatubular casing |22, which delivers the compressed air to the intake of amain compressor I|2 towards the rear end of the turbine shaft. Thiscompressor feeds air through a manifold |28 into the breech ends of thecombustion chambers II3 which are spaced in an annular series around thecasing |22; a

Vseparately acting compressor |25 formed by the rear face of the maincompressor delivers a supply of air through fuel injectors |03 into thebreech ends of the combustion chambers, this fuel-injection air beingdrawn partly from the interior of the hollow shaft IIO by way of holes|21 near its rear end and to a much larger extent from the exterior, ashereinafter described, while the fuel is supplied by the auxiliaryequipment |04 driven by the rear end of the turbine shaft. It will benoted that the rear compressor |25 feeds the combustion chamber burnersdirectly with air conveying fuel from the injectors |03, while the maincompressor I|2 furnishes air through the manifold |28 at higher speed,at higher pressure and in greater quantity to maintain the names in fullactivity.

The hot gases. discharged from the combustion chambers I I3 through thestator guide vanes I I4 impinge upon the turbine blades ||5 at theperiphery of the rotor III; these blades are preferably arranged formulti-stage expansion of the gases, intermediate guide vanes |I4a beingprovided between the stages, and the blades II5, II5 of the successivestages being of increasing height and external diameter to `allow forthe expansion of the gases, while at the same time increasing thecentrifugal effect. After passing the final-stage blades I I5a, thegases are reversed through substantially 180 degrees without shocks oreddies, but taking advantage of the outward diifusion or spread of thegases under centrifugal effect, by means of a curved deflecting vsurface|05; the latter, which serves asa reaction Ydevice to assist thepropulsion of the aircraft, directs the gases rearwards over and betweenthe combustion chambers I I3, within the progressively expanding annularspace between an outer cylindrical casing |06 and a slightly convergentpartition |01 through which the combustion chambers extend, so that thegases flow along the gaps between the said chambers, the rear endof thispartition being suitably apertured to allow passage of the compressedair to the injectors |03 and manifold |28. The said injectors andmanifold are provided with external fins |08, in addition to internaldeectors |29, the whole of these fittings being preferably castintegrally in high-conductivity metal, so that a considerable andintensive heating of the compressed air is effected atthis point by thehighspeed gases returning over and between the combustion chambers H3,outside the partition |01. Such a heat-exchanger, with its externalsurfaces bathed or swept by very .hot gases travelling at high speed,procures a much .more eflicient transfer of heat through the Walls ofthe injectors and manifold than would their immersion in a hot gaseousmedium of low speed or even stagnant.

From the moment when the hot gases have passed the finned injectors |03fed by the rear compressor |25, they are directed into a jet device ofthe well-known kind, comprising two coaxial cones |09, I09a with a freespace between them leading to a jet outlet for the rearward discharge ofthe gases at high speed; the outer cone |09 forms a continuation of thecasing |06, while the inner cone |0911 forms a `continuation of thepartition |07, this inner cone or bullet enclosing the auxiliaryequipment |04. The inner cone is stayed in relation to the outer cone byflared or stream-line Webs |26, suitably heat-insulated, but instead ofthese webs being solid or closed at the ends, they are made hollow (asseen in section in the case of the uppermost one) to connect withopenings through the two cones so that the external air can pass freelyand substantially unheated to the intake of the rear compressor |25. Theopenings through the Webs |26 may be made of sufficient size to allowaccess to the auxiliary equipment for the purpose of inspection andmaintenance.

As in the case of the finned injectors |63 and manifold |28, the rotorIII may be `cast in tegrally with its rim |33, turbine blades |I5 andcompressor vanes I`I5, the whole being made of a high-conductivity metalto facilitate the transmission of heat from the blades to the vaneswhich are cooled by the air admitted through the rotor. The rim |33 fitsinto grooves |41 on the adjacent edges of the stator walls, providing africtionless seal against leakage of the turbine gases to the interiorof the rotor.

Due to the fact that the rearward flow of the exhaust gases from thedeflecting surface |05, which itself produces a tractive force greaterthan the static thrust of the jet, takes place mainly in the gapsbetween the spaced combustion chambers `||3, the diameter of the outercasing |06 need be very little greater than the overall diameter ofthese chambers; consequentlyI the improved turbine presents theimportant advantage of relatively small external diameter, withresulting low weight, in addition to the advantages of superior coolingof the moving parts due to the admission of cold air through theapertured rotor, and that of fuel economy due primarily to thecombination of a turbine-driven airscreW with the propulsive jet, andsecondly to the preheating of the compressed air which is controlled bythe design of the fins |08 and by the provision of heat-insulation atsuitable points.

Fig. 10 illustrates the arrangement of a gas turbine according to thepresent invention as applied to the propulsion of an aircraft by a jetdevice 245 without the provision of an airscrew driven by the turbineshaft; in this case, the shaft 2|0 has the turbine rotor 2|| mountedtowards the rear end, and the double compressor 2|2, 225 mounted towardsthe front end, with the guide vanes 2 |8 and the multiple rotors 220 andstators 22| of the axial-type compressor disposed along the shaft. Thecombustion chambers 2 |3 are arranged in an annular series around theturbine, their rear ends being tted with guide vanes 2|4 delivering thehot gases upon the turbine blades 2 5 integral with the rim 233 of therotor; the exhaust gases are discharged to the jet device 245 by way ofthe free space between two coaxial cones 209, 209a.

The double compressor 2|2, 225 driven by the turbine shaft feeds thecombustion chambers 2|3 in a similar manner to that described withreference to Fig. 9, the fuel being supplied by the auxiliary equipment204 and the air for the compressor 225 being admitted partly through thehollow shaft by way of holes 221 and partly from the exterior throughthe intake openings 226. The main supply of air to the compressor 2|2 isadmitted through the apertured rotor 2|| having vanes 2 I6 which form aprimary element of the axial-type compressor; since, however, this rotor2|'| is arranged at the rear end of the turbine, where the freeadmission of air is hindered by the inner cone or bullet 209a, the wholeof the turbine is encased by a power egg or casing 206 having itsforward end 201 left open and slightly flared so that the air enters byram effect. The air passes over and between the injectors 203, manifold228 and combustion chambers 2|3, all of which are suitablyheat-insulated to prevent loss of heat to the incoming air; as the airtravels along the interior of the casing 206, it moves clear of thecombustion chambers 2|3 and flows outside the cone 209 at its junctionwith the turbine stator 205, beyond which the casing 206 is reduced indiameter, its rear end being suitably connected to the outer cone 209. Anumber of air ducts or inlets 230, of stream-line section, are carriedthrough the cones 209, 209a, between the stator 205 and the rear end ofthe casing 206, the walls of these ducts acting to hold the two cones inrelative position, while the passage ways through the ducts enable theair from the rear end of the casing 206 to pass inwards to the interiorof the inner cone 209a and thus to reach the apertured rotor 2| with aconsiderable initial pressure due to the ram effect at the intake end201 of the casing.

It will be noted that both air intakes are located at the normal orforward end of the turbine, where the auxiliary equipment 204 isinstalled, and that the incoming air is maintained substantiallyunheated until it enters the compressors, heat-insulation being providedwhere necessary to prevent transfer to heat. The jet outlet 245 at therear end may be arranged for a slight expansion of the gases under theeffect of centrifugal force, by giving the outlet a slightly divergentshape, as indicated at 246. As in the previous examples of construction,the turbine rotor has its annular rim 233 fitting into grooves 241 onthe adjacent edges of the stator walls, thus providing a frictionlessseal to prevent leakage of the high-velocity gases from the combustionchambers into the interior of the rotor 2| I.

What I claim is:

l. A gas turbine comprising a rotor having peripheral blades, a hollowshaft driven by said rotor, a casing around said shaft, at least onecombustion chamber furnishing hot gases, means for applying said gasesto said blades in a direction parallel to said shaft for driving saidrotor, a main compressor driven by said shaft for compressing air tofeed said combustion chamber, said rotor and said main compressor beingarranged towards opposite ends of said casing, and an axial-typecompressor with multiple rotors and stators inside said casing, saidfirst-mentioned rotor being apertured to admit cold air to saidaxial-type compressor, and said axial-type compressor forcing the air tosaid main compressor in a direction opposite to the ow of said gasespast said blades, in combination with another compressor driven by saidshaft and drawing cold air through said hollow shaft, said othercompressor feeding air to at least one combustion chamber furnishing hotgases to drive said firstmentioned rotor.

2. A gas turbine comprising a rotor having peripheral blades, a hollowshaft driven by said rotor, a casing around said shaft, at least onecombustion chamber furnishing hot gases, means for applying said gasesto said blades in a direction parallel to said shaft for driving saidrotor, a main compressor driven by said shaft for compressing air tofeed said combustion chamber, said rotor and said main compressor beingarranged towards opposite ends of said casing, said combustion chamberbeing located between said compressor and said blades, and an axial-typecompressor with multiple rotors and stators inside said casing, saidfirst-mentioned rotor being apertured to admit cold air to saidaxial-type compressor, and said axial-type compressor forcing the air tosaid main compressor in a direction opposite to the flow of said gasespast said blades, in combination with another compressor driven by saidshaft and drawing coldair through said hollow shaft, said othercompressor being nonsymmetrical in relation to said main compressor,said axial-type and main compressors operating in series, and said othercompressor operating in parallel to said series-operating axial-type andmain compressors.

3. A gas turbine for the propulsion of vehicles including aircraft,comprising a rotor having peripheral blades, a shaft driven by saidrotor, a plurality of combustion chambers furnishing hot gases, meansfor applying said gases to said blades for driving said rotor, anaxial-type compressor and two centrifugal compressors, all saidcompressors being driven by said shaft for compressing air to feed saidcombustion chambers, said rotor being arranged towards the rear end ofsaid shaft, said centrifugal compressors being arranged towards theforward end of said shaft.

9. saidrotor being apertured for passage of cold air in a' forwarddirection to said axial-type compressor; a manifold of annular formcollecting the-exhaust'` gases flowing in a rearward direction from saidrotor blades, and at least one outlet from said manifold for dischargingthe exhaust gases rearwardly of the vehicle.

4. A gas turbine for the propulsion of vehicles including aircraft,comprising a rotor having peripheral blades, a hollow shaft driven bysaid rotor; a plurality of combustion chambers furnishing hot gases,means for applying said gases tofsaidblades for driving said rotor, anaxial-type compressor, ak main centrifugal compressor and anothercentrifugal compressor, all said compressors being driven by said shaftfor compressing air to feed said combustion chambers, part f saidair'being drawn through said hollow shaft, said rotor being arrangedtowards the rear end of said shaft, said centrifugal compressors beingarranged towards the forward end of said shaft, said rotor beingapertured for passage of cold air in aV forward direction to saidaxial-type compressor, a manifold of annular form collecting the exhaustgases fiowing in a rearward direction from said rotor blades, saidmanifold having two semiannular branches of spiral shape, two outletsfrom said branches, and means for combining the gases'discharged throughsaid two outlets to deliver a single rearwardly directed jet for thepropulsion' of the vehicle.

5. Av gas turbine for the propulsion of vehicles including aircraft,comprising a rotor having peripheral blades, a shaft driven by saidrotor, a plurality of combustion chambers furnishing hot gases', meansfor applying said gases to said blades for driving said rotor, anaxial-type compressor and two centrifugal compressors, all saidcompressors being driven by said shaft for compressing air to feed saidcombustion chambers, said rotor being arranged towards the forward endof said shaft, said centrifugal compressors being arranged towards therear end of said shaft, said rotor being apertured for passage of coldair in a rearward direction by ram effect to the intake of saidaxial-type compressor, a curved defiecting surface for reversing theforwardly exhausting gases from said rotor blades and directing saidgases rearwardly between said combustion chambers, and means forcollecting said rearwardly directed gases to deliver a rearward jet forthe propulsion of the vehicle.

6. A gas turbine for the propulsion of vehicles including aircraft,comprising a rotor having peripheral blades, a shaft driven by saidrotor, a plurality of combustion chambers furnishing hot gases,A meansfor applying said gases to said blades for driving said rotor, anaxial-type compressor and two centrifugal compressors, all saidcompressors being driven by said shaft for compressing air to feed saidcombustion chambers, said rotor being arranged towards the forward endof said shaft, said centrifugal compressors, being arranged towards therear end of said shaft, said rotor being apertured for passage of coldair in a rearward direction by ram effect to the intake of saidaxial-type compressor, a curved deflecting surface for reversing theforwardly exhausting gases from said rotor blades and directing saidgases rearwardly past said combustion chambers, means for effectingheat-exchange between said rearwardly directed gases and the air fed tosaid combustion chambers by said centrifugal compressors, and means forcollecting said rearwardly directed gases to deliver a rearward jet forthe propulsion of the vehicle.

7. A gas turbine for the propulsion of4 aircraft, comprising a rotorhaving peripheral blades arranged for at least two stages of gasexpansion, a hollow shaft driven by said rotor, a plurality of spacedcombustion chambers furnishing hot gases, means for applying said gasesto said blades for driving said rotor, an axial-type compressor and twocentrifugal compressors driven by said shaft for compressing air to feedsaid combustion chambers, part of said air being drawn through saidhollow shaft, said rotor being arranged towards the forward end of saidshaft, said centrifugal compressors being arranged towards the rear endof said shaft, an airscrew at the forward end of said shaft, saidairscrew being driven by said' shaft, said rotor being apertured forpassage of cold air from said airscrew in a rearward direction to theintake of said axial-type compressor, a curved deiiecting surfacereceiving the forwardly exhausting gases from said rotor blades, saiddeflecting surface being adapted to reverse the flow of said gases bytaking advantage of their outward diffusion under centrifugal eiect andto direct them rearwardly around said combustion chambers, an outercasing around said combustion chambers, said casing being adapted toenclose said rearwardly directed gases and to force them into the gapsbetween said spaced combustion chambers, and means for collecting thegases rearwards of said combustion chambers to deliver a single jet forthe propulsion of the aircraft.

8. A gas turbine for the propulsion of aircraft, comprising a rotorhaving peripheral blades, a hollow shaft driven by said rotor, aplurality of combustion chambers arranged in an annular series aroundsaid shaft, means for applying hot gases from said combustion chambersto said blades in a direction parallel to said shaft for driving saidrotor, means for collecting the gases in rear of said blades to delivera jet for the propulsion of the aircraft, an axial-type compressor andtwo centrifugal compressors, all said compressors being driven by saidshaft for compressing air to feed said combustion chambers, part of saidair being drawn through said hollow shaft by one of said centrifugalcompressors, said rotor being arranged towards the rear end of saidshaft, said centrifugal compressors being arranged towards the forwardend of said shaft, said rotor being apertured for passage of air in aforward direction to the intake of said axial-type compressor, an outercasing around said combustion chambers, the forward end of said casingbeing open for admission of cold air by ram effect, said air flowingalong the gaps between said combustion chambers, and passages leadingfrom the rear end of said casing to convey said air to said aperturedrotor, said passages extending transversely through said collectingmeans.

9. A gas turbine comprising a rotary shaft, an axial-type compressorwith multiple rotors and stators, the rotors of said compressor beingdriven by said shaft, a stator element at the inlet end of saidcompressor, a rotor element adjacent to said stator element, said rotorelement including turbine blades upon its periphery and helicoidalspokes secured to said shaft, two centrifugal compressors driven by saidshaft and located at the other end of said axial-type compressor, saidhelicoidal spokes co-operating with said stator element to force airinto said axial-type compressor for feeding one of said centrifugalcompressors, combustion chambers arranged to re-l ceive air from saidcentrifugal compressors and to furnish hot gases of combustion, andmeans for applying said hot gases to said turbine blades for drivingsaid rotor element and shaft, the direction of said hot gases inrelation to said rotor element being opposite to the direction of saidair in relation thereto.

10. A gas turbine comprising a rotary shaft, an axial-type compressorwith multiple rotors and stators, the rotors of said compressor beingdriven by said shaft, a stator element at the inlet end of saidcompressor, a rotor element adjacent to said stator element, said rotorelement including peripheral turbine blades integral with helicoidalspokes secured to said shaft, two centrifugal compressors driven by saidshaft and located at the other end of said axial-type compressor, saidhelicoidal spokes co-operating with said stator element to force airinto said axial-type compressor for feeding one of said centrifugalcompressors, combustion chambers arranged to receive air from saidcentrifugal compressors and to furnish hot gases of combustion, andmeans for applying said hot gases to said turbine blades for drivingsaid rotor element and shaft, the direction of said hot gases inrelation to said rotor element being opposite to the direction of saidair in relation thereto.

l1. A gas turbine comprising a rotary shaft, an axial-type compressorwith multiple rotors and stators, the rotors of said compressor beingdriven by said shaft, a stator element at the inlet end of saidcompressor, a rotor element adjacent to said stator element, said rotorelement including turbine blades upon its periphery and helicoidalspokes secured to said shaft, two centrifugal compressors driven by saidshaft and located at the other end of said axial-type compressor, saidhelicoidal spokes co-operating with said stator element to force airinto said axial-type compressor for feeding one of said centrifugalcompressors, combustion chambers arranged to receive air from saidcentrifugal compressors and to furnish hot gases of combustion, andmeans for applying said hot gases to said turbine blades for drivingsaid rotor element and shaft, said gas-applying means including bafflesfor preventing leakage of said hot gases adjacent to said rotor elementinto the air forced into said axial-type compressor,

and the direction of said hot gases in relation to y said rotor elementbeing opposite to the direction of said air in relation thereto.

l2. A gas turbine comprising a hollow rotary shaft, an axial-typecompressor with multiple rotors'and stators, the rotors of saidcompressor i being driven by said hollow shaft, a stator element at theinlet end of said compressor, a rotor element adjacent to said statorelement, said rotor element including turbine blades upon its peripheryand helicoidal spokes secured to said shaft, two centrifugal compressorsdriven by said shaft and located at the other end of said axial-typecompressor, one of said centrifugal compressors drawing air through oneend of said hollow shaft, the other of said centrifugal compressorsreceiving air from said axial-type compressor, said helicoidal spokesco-operating with said stator element to force air into said axial-typecompressor, combustion chambers arranged to receive air from saidcentrifugal compressors and to furnish hot gases of combustion, andmeans for applying said hot gases to said turbine blades for drivingsaid rotor element and shaft, the direction of said hot gases inrelation to said rotor element being opposite to the direction of saidair in relation thereto.

13. A gas turbine comprising a hollow rotary shaft, an axial-typecompressor with multiple rotors and stators, the rotors of saidcompressor being driven by said hollow shaft, a stator element at theinlet end of said compressor, a rotor element adjacent to said statorelement, said rotor element including turbine blades upon its peripheryand helicoidal spokes secured to said shaft, a centrifugal compressordriven by said shaft and located at the other end of said axialtypecompressor, said helicoidal spokes co-operating with said stator elementto force air into said axial-type compressor for feeding saidcentrifugal compressor, another centrifugal compressor driven by saidshaft and located back-to-back of said first-mentioned centrifugalcompressor, said other centrifugal compressor drawing air through oneend of said hollow shaft, at least one combustion chamber arranged toreceive air from said centrifugal compressors and to furnish hot gasesof combustion, and means for applying said hot gases to said turbineblades for driving said rotor element and shaft, the direction of saidhot gases in relation to said rotor element being opposite to thedirection of said air in relation thereto.

JULES ANDR NORBERT GALLIOT.

REFERENCES CITED The following references are of record in the le ofthis patent:

UNITED STATES PATENTS Number Name Date 2,080,425 Lysholm May 18, 19372,224,260 Galliot Dec. 10, 1940 2,326,072 Seippel Aug. 3, 1943 2,396,068Youngash Mar. 5, 1946 2,405,164 Pavlecka Aug. 6, 1946 2,410,804 BaumannNov. 12, 1946 2,423,183 Forsyth July l, 1947 2,428,330 Heppner Sept. 30,1947 2,455,458 Whittle Dec. 7, 1948 2,477,798 Griffith Aug. 2, 1949

